A Sun acquisition phase is for example carried out by a spacecraft, such as a satellite, after separation from a launcher of said satellite. Such a Sun acquisition phase is referred to in the rest of the description as an “initial Sun acquisition phase”.
Specifically, after separation from the launcher, the satellite is autonomous and needs to ensure its electrical autonomy by trying to point its solar generators toward the Sun in order to provide electrical power to a platform of said satellite and in order to recharge the batteries of said satellite. Conventionally, the batteries are charged before the satellite is released, and, after separation and before the solar generators are deployed and are directed toward the Sun, the electrical power necessary for the operation of the platform of the satellite is provided by said batteries.
This initial Sun acquisition phase is made difficult by the fact that, at the time of separation from the launcher, the satellite may be released with a high rotational speed: typically up to 3°/s in an LEO orbit (“Low Earth Orbit”) and of the order of 1°/s in a GEO orbit (“Geostationary Orbit”). The initial Sun acquisition phase therefore comprises a step of reducing the rotational speed of the satellite. This reduction of the rotational speed of the satellite needs to be short so as not to risk fully discharging the batteries of said satellite.
Thus, the satellite comprises actuators used in order to reduce the angular momentum of the satellite after separation from the launcher.
It is in particular known, for satellites in an LEO orbit, to employ magnetic torquers using the Earth's magnetic field in order to form torques capable of reducing the angular momentum of the satellite. For satellites in an MEO (“Medium Earth Orbit”) or GEO (or GTO—“Geostationary Transfer Orbit”) orbit, for which the Earth's magnetic field is negligible, chemical thrusters are generally used. Such chemical thrusters are advantageous in that they can be used immediately after separation from the launcher, in that they do not require a high electrical power, and in that they can apply large torques to the satellite, which makes it possible to reduce the angular momentum and to place the satellite in the acquisition orientation in a very short time (of the order of a few minutes).
After the angular momentum of the satellite has been reduced sufficiently and the satellite has been placed in the acquisition orientation, the solar generators are deployed and directed toward the Sun in order to supply the platform of said satellite and recharge the batteries.
It is currently envisioned that future satellites will no longer be equipped with chemical thrusters, but only with electrical (plasma) thrusters. The initial Sun acquisition phase would then become problematic, particularly for satellites in an MEO/GEO/GTO orbit, because electrical thrusters cannot replace chemical thrusters for reduction of the angular momentum of the satellite after separation.
This is because electrical thrusters cannot be used during the first days following separation from the launcher. Furthermore, the thrusters require a high electrical power (of the order of a few kW) and can only apply low torques to the satellite (of the order of 0.2 N·m). Consequently, reducing the angular momentum of the satellite by means of electrical thrusters would require a high electrical power for a long time, which is incompatible with the batteries installed in current satellites, which have a capacity of the order of 10 kW·h. Such a capacity makes it possible to supply the electrical thrusters for only about one hour, which is insufficient to reduce the angular momentum of the satellite sufficiently in view of the torques which can be formed by current electrical thrusters.
Current satellites are generally equipped with electrical inertial actuators, such as flywheels (reaction wheels, gyroscopic actuators), which are used in order to control the attitude of said satellite along three axes. Such inertial actuators generally have an insufficient capacity to compensate on their own for the angular momentum of the satellite after separation, but may be used after the other actuators (magnetic torquers, chemical thrusters, etc.) have reduced said angular momentum of said satellite sufficiently.
One alternative might consist in dimensioning the electrical inertial actuators so that they have a capacity to compensate on their own for the angular momentum of the satellite after separation. This, however, would lead to a significant increase of the mass and the volume of said inertial actuators. By way of example, the initial angular momentum of a satellite after release may be of the order of from 500 N·m·s to 1000 N·m·s, while the capacity of the inertial actuators installed in a satellite is generally of the order of from 50 N·m·s to 100 N·m·s. Furthermore, even though the inertial actuators might be dimensioned so as to have a capacity of 1000 N·m·s, particularly unfavorable separation conditions could still transfer an initial angular momentum of more than 1000 N·m·s to the satellite, which is beyond the capacity of the inertial actuators.
It should be noted that the Sun acquisition phase may also be carried out when the satellite is in survival mode. The comments above remain valid for such a Sun acquisition phase in survival mode.